Virtually all modern commercial jet aircraft in use today utilize a single wing extending laterally in both directions from a central portion of the fuselage. Such aircraft are designed and loaded so that the overall aircraft center of gravity will be located just forward of the aerodynamic center of lift of the wing; as required by stability considerations.
In flight the aircraft is balanced about the pitch axis by a balancing tail load, which normally acts downwardly. The balancing tail load is provided by a stabilizer, or horizontal tail surface, at the aft end of the fuselage. The stabilizer has small control surfaces called elevators to adjust the magnitude and direction of the tail load. The lift provided by the wing must be sufficient not only to lift the gross weight of the aircraft, but to compensate for any downward acting tail load. Accordingly, when such tail loads are high, as at takeoff, the allowable gross weight of an aircraft is significantly reduced. Moreover, drag loads are high on both the wing and tail surfaces.
In such conventional single wing aircraft, the fuselage structure is effectively two cantilever beams; one extending forward and the other aft of the wing. Maximum fuselage bending moments therefore occur near the wing in the central region of the fuselage. Furthermore, as will be more fully developed in the discussion which follows, because of "area-ruling" requirements, the fuselage structure of high speed aircraft must normally be reduced in diameter in this same central region of the fuselage. Such reduction in diameter of the fuselage is in widespread use today in supersonic aircraft design, and is known in the industry as "coke-bottling" of the fuselage. Not only are substantial structural weight penalties incurred by having the maximum loads at the minimum section, but in commercial passenger aircraft undesirable seat and aisle arrangements are required, and main landing gear stowage space is severely limited, sometimes requiring external fairing.
In such conventional single wing aircraft, the wing structure is effectively two cantilever beams; one beam extending in each direction laterally from the fuselage. Design bending moments increase rapidly from a minimum at the tip to a maximum at the root of each wing semi-span. In order to reduce drag, vertically thin and higher aspect ratio (wing span/effective average wing chord length) wings are desirable. However, such longer wings tend to create prohibitively high structural bending moments on the desired thin wing sections. Considerable research has been done in recent years on optimum wing cross-sectional shapes. "Supercritical airfoil" and "natural laminar flow" wing sections have been developed to reduce shockwave formation and promote laminar or non-turbulent flow over the section. Such wing shapes tend to have aerodynamic centers located farther aft on their sections (of the order of 40% chord as opposed to some 30% chord for a typical conventional section). This further aggrevates the balancing tail load problem previously discussed, requiring that still more downward acting tail load be made available.
Roll control in a conventional single wing aircraft is normally achieved by ailerons located in outboard regions of the wing. Lift control, as for takeoff and landing, is achieved by flaps extending from aft (and sometimes forward) portions of inboard regions of the wing. Spoilers extending upwardly from the wing are often used as speed brakes or to reduce lift. Accordingly, the wing, particularly its trailing edge, is crowded with primary and secondary control surfaces. The horizontal tail, with its elevator control surface, is normally dedicated exclusively to pitch control. The fin, or vertical tail, and its rudder are normally dedicated to yaw control.
Because of the complexity of modern aircraft, and the number of control surfaces, they are becoming increasingly difficult to fly manually. Accordingly, complex computerized "stability augmentation" and automatic pilot systems have been designed to automatically adjust control surfaces and reduce pilot work load. For safety reasons, it is extremely desirable that these computerized systems be redundant; i.e., that more than one, and preferably up to four, of such systems be completely independent, and each capable of achieving a desired maneuver for the aircraft. However, such multiple redundancy is extremely difficult, if not impossible, to attain in conventional single wing designs. The space available for the multitude of control surfaces required for such redundancy is inherently limited, and only one region of the aircraft is available in some cases to achieve the desired degree of control (e.g., stabilizer for pitch, and fin for yaw control).
Accordingly, it is a general objective of this invention to provide an aircraft configuration in which: (1) a downward acting balancing tail load is not required; (2) it is not necessary to reduce the fuselage diameter in its central region when area-ruling is required; (3) structural weight of the fuselage is reduced; (4) the wing structural weight is reduced; (5) higher aspect ratio wing sections may readily be utilized for reduced drag; (6) recently developed supercritical and natural laminar flow wing sections may be used; and (7) there is a plurality of flight control systems, each independently able to control the aircraft to provide redundancy for flight safety.